PDF/ocr_dataset_1023.jsonl

{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 36, "total_pages": 47, "image_filename": "19930093773_p36.jpg", "text": "NACA RM E9G09\n35\n\n1159\n\n| Altitude (ft) | |\n| :--- | :--- |\n| $\\circ$ | 5,000 |\n| $\\diamond$ | 15,000 |\n| $\\triangle$ | 25,000 |\n| $\\square$ | 35,000 |\n| $\\nabla$ | 45,000 |\n| $\\triangledown$ | 50,000 |\n\nCorrected net thrust, $F_n/\\delta$, lb\n\nCorrected engine speed, $N/\\sqrt{\\theta}$, rpm\n\n[NACA logo]\n\n(a) Net thrust.\n\nFigure 6. - Effect of altitude on variation of corrected engine performance with corrected engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:54:58.573411+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 7, "total_pages": 28, "image_filename": "19930085471_p7.jpg", "text": "NACA RM No. L8J11\nCONFIDENTIAL\n5\n\nRESULTS AND DISCUSSION\n\nThe significant flutter parameters and the results of the calculations are given in table I. As indicated by test methods, a large number of runs were made on wings which did not flutter. Table I gives only the results for the wings for which flutter was observed. In altering the model to obtain flutter, the center of gravity was moved back in steps of about 2 percent of the chord. Consequently, the position of the center of gravity which would produce flutter is known to about 2 percent of the chord and lies between the position that did not lead to flutter and the position that produced flutter. The results are sensitive to center-of-gravity position and this may account for some of the scatter of the data. Contributing also to the scatter of the data are the inaccuracies in obtaining the wing parameters, the effect of the degree of penetration of the model into the tunnel, and errors in the determination of the flutter-speed coefficients arising from the method of interpolation.\n\nSome of the results listed in table I are presented in figures 6 and 7. In figure 6 the theoretical and experimental flutter-speed coefficients are compared. The fact that the experimental data fall above the $45^\\circ$ line, for the most part, indicates that the theory of reference 1 is generally conservative as far as application to cantilever wings is concerned. Considering that a slight inaccuracy in the location of the center of gravity has a large effect and considering also unaccounted effects of section shape, aspect ratio, and Mach cone, the agreement is not unsatisfactory. The theoretical and experimental flutter frequencies are compared in figure 7; the experimental frequencies ranged from about 0.61 to 1.09 times the theoretical values. In all cases the mode at flutter appeared to consist mainly of a coupling of first bending and first torsion modes.\n\nSince the present investigation is of a preliminary nature and covers a wide range of parameters, no attempt was made to isolate the effects of separate parameters such as the mass-density parameter, frequency ratio, elastic axis, and center of gravity which are treated by the two-dimensional theory, or parameters such as aspect ratio not treated by the theory.\n\nAn attempt was made to investigate some of the possible effects of airfoil section shape on flutter. The intermingling of the data for the various models (figs. 6 and 7) suggests that the section shape has no very pronounced effect on flutter at Mach number 1.3. However, more difficulty due to divergence was experienced with thick models having blunt leading edges. This is in accord with the higher-order method of calculation (order higher than in the linear method) for pressure distribution at supersonic speeds in steady flow which shows that the center of pressure may move ahead of the 50-percent-chord position for thick blunt-nosed airfoils particularly at Mach numbers near unity. It was\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T04:55:00.952624+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 24, "total_pages": 99, "image_filename": "19930082511_p24.jpg", "text": "$$\n\\left. \\begin{array} { c } { { \\tt A = 0 } } \\\\ { { \\tt B = - \\frac { \\Gamma ^ { \\prime } } { 2 \\sqrt { 1 + y _ { 1 } ^ { 2 } } } } } \\end{array} \\right\\}\n$$\n\nIf the point of evaluation is also on the axis (that is, $\\xi = \\xi + \\frac { 1 } { 2 }$ , whence $z = \\mathrm { i } y$ ), then\n\n$$\n{ \\mathfrak { q } } _ { 1 } \\left( \\xi + { \\frac { 1 } { 2 } } , \\xi _ { 1 } + { \\frac { 1 } { 2 } } \\right) = - 1 { \\Gamma } ^ { \\prime } \\ { \\frac { { \\mathfrak { y } } { \\mathfrak { y } } _ { 1 } } { { \\mathfrak { y } } ^ { 2 } - { \\mathfrak { y } } _ { 1 } ^ { 2 } } } \\ { \\sqrt { \\frac { { \\mathfrak { y } } ^ { 2 } + 1 } { { \\mathfrak { y } } _ { 1 } ^ { 2 } + 1 } } } + { \\frac { { \\mathrm { i } } { \\Gamma } ^ { \\prime } } { 2 { \\pi } \\left( \\xi - \\xi _ { 1 } \\right) } }\n$$\n\nThus, if the vortex is on the axis, the interference velocity at all points on the axis has only a vertical component.\n\nThe interference velocity at the vortex itself is the limit of expression (5) as z approaches $z _ { 1 }$ . The term containing $z - \\bar { z } _ { 1 }$ offers no difficulties and its limit is readily evaluated:\n\n$$\n\\operatorname* { l i m } _ { z \\to z _ { 1 } } - \\frac { 1 } { z - \\bar { z } _ { 1 } } ( \\tt A + B z ) \\sqrt { 1 - z ^ { 2 } } = - \\frac { 1 } { 2 \\bar { y } _ { 1 } } ( \\tt A + B z _ { 1 } ) \\sqrt { 1 - z _ { 1 } ^ { 2 } } = \\frac { \\Gamma ^ { \\prime } z _ { 1 } } { 4 y _ { 1 } }\n$$\n\nwhere the last equality follows from equation (4). The remainder of equation (5), after $\\pi ( \\xi - \\xi _ { 1 } )$ is replaced by $\\log { \\frac { z } { z _ { 1 } } }$ , is\n\n$$\n\\operatorname* { l i m } _ { z \\to z _ { 1 } } \\left[ { \\frac { ( \\mathbb { A } + \\mathbb { B } z ) { \\sqrt { 1 - z ^ { 2 } } } } { z - z _ { 1 } } } + { \\frac { \\Gamma ^ { \\prime } } { 2 \\log { \\frac { z } { z _ { 1 } } } } } \\right] = \\operatorname* { l i m } _ { z \\to z _ { 1 } } { \\frac { ( \\mathbb { A } + \\mathbb { B } z ) { \\sqrt { 1 - z ^ { 2 } } } \\log { \\frac { z } { z _ { 1 } } } + { \\frac { \\Gamma ^ { \\prime } } { 2 } } ( z - z _ { 1 } ) } { ( z - z _ { 1 } ) \\log { \\frac { z } { z _ { 1 } } } } }\n$$\n\nThis expression is of the form $\\frac { 0 } { 0 }$ for $z = z _ { 1 }$ . Differentiating numerator and denominator, according to L'Hôspital's rule still leaves both equal to zero at $z = z _ { 1 }$ (that the derivative of the denominator is zero at $z = z _ { 1 }$ is obvious; that the derivative of the numerator is also zero at $z = z _ { 1 }$ can be verified with the aid of equation (4)). A second differentiation yields the following expression for the limit as z approaches $z _ { 1 }$ :", "timestamp": "2026-07-22T04:55:05.708673+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 15, "total_pages": 58, "image_filename": "19930082617_p15.jpg", "text": "14\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 28.8 X 10^3\n2 57.6 X 10^3\n3 86.4 X 10^3\n4 115.2 X 10^3\n5 144.0 X 10^3\n6 201.6 X 10^3\n\n3.86\"\n[Figure: Diagram showing Band H with section A-A]\n\nDistance from horizontal diameter, in.\nStrain\n\n[Graph showing strain distribution with curves labeled 1, 2, 3, 4, 5, 6]\n\nNACA\n\nFigure 3.- Strain diagram of cylinder 72. Band H.", "timestamp": "2026-07-22T04:55:06.445693+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 14, "total_pages": 36, "image_filename": "19930082614_p14.jpg", "text": "12\nNACA TN 1939\n\nhigh-speed airplane because a large part of the latter maneuver would\nconsist of the dive entry and pull-out. The effects of curvature of\nthe flight path can be evaluated by an extension of the graphical\nmethod just described.\n\nWhen the airplane is following a path of increasing or decreasing\ninclination, a calculation of the changes in altitude is complicated by\nthe necessity of extrapolating along the curved path. One means of\nmaking such an extrapolation is to assume that the airplane follows a\npath consisting of a series of circular arcs. The curvature of the\nflight path, which is a function of the forces normal to the direction\nof motion, is evident to the pilot as a normal acceleration. In order\nto permit the changes in altitude to be expressed as a function of the\nnormal acceleration felt by the pilot, motion in a vertical plane is\nfirst considered. The altitude change during the period of time $\\Delta t$\ncan be computed by the relation\n\n$$\\Delta h = r [\\cos \\gamma_1 - \\cos (\\gamma_1 + \\Delta \\gamma)] \\quad (8)$$\n\nwhere $r$ is the radius of the arc of the flight path, $\\gamma_1$ is the\nflight-path angle at the beginning of the period, and $\\Delta \\gamma$ is the\nangular change in the direction of motion during the time $\\Delta t$.\n\nThe actual acceleration normal to the direction of flight is equal\nto the indicated normal acceleration minus the normal acceleration due\nto the weight of the airplane. If $n$ is the indicated normal accelera-\ntion factor,\n\n$$\\frac{V^2}{r} = g (n - \\cos \\gamma) \\quad (9)$$\n\nIn figure 5(a) the acceleration factor $n$ is plotted against the\nquantity $(n-\\cos \\bar{\\gamma})$ for values of $\\bar{\\gamma}$ from $0^\\circ$ to $90^\\circ$. The curves for\nnegative values of $\\bar{\\gamma}$ are the same as for positive values.\nFigure 5(b) presents the variation of airspeed with $(n-\\cos \\bar{\\gamma})$ for\nconstant radii of flight-path curvature. Example guide lines show how\nthe radius may be found by projecting the abscissa from a point in\nfigure 5(a) to the line for $\\cos \\bar{\\gamma} = 0$ ($\\bar{\\gamma} = 90^\\circ$, positive $n$) and con-\ntinuing at this ordinate to the desired value of average airspeed in\npart (b).\n\nFigures 5(c) and 5(d) give a graphical solution for $\\Delta \\gamma$ as a\nfunction of $(n - \\cos \\bar{\\gamma})$, $\\bar{V}$, and $\\Delta t$. The graphs were obtained from\nthe following equations. Since $r$ is considered to be constant during\nthe time interval $\\Delta t$,", "timestamp": "2026-07-22T04:55:07.374157+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 10, "total_pages": 14, "image_filename": "19930082712_p10.jpg", "text": "8\nNACA TN 1998\n\n4. For the smooth condition, the NACA 8-H-12 airfoil had a lower\nmaximum lift coefficient than either the NACA 23012 or the NACA 0012\nsections at comparable Reynolds numbers. The addition of leading-edge\nroughness, however, affected the NACA 8-H-12 airfoil less adversely\nthan the other two sections. The drag coefficient of the NACA 8-H-12\nairfoil measured at the design lift was, in general, lower than that of\nthe NACA 0012 and the NACA 23012 sections for both surface conditions.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., November 8, 1949\n\nREFERENCES\n\n1. Tetervin, Neal: Tests in the NACA Two-Dimensional Low-Turbulence\nTunnel of Airfoil Sections Designed to Have Small Pitching\nMoments and High Lift-Drag Ratios. NACA CB 3L13, 1943.\n\n2. Stivers, Louis S., Jr., and Rice, Fred J., Jr.: Aerodynamic\nCharacteristics of Four NACA Airfoil Sections Designed for\nHelicopter Rotor Blades. NACA RB L5K02, 1946.\n\n3. Schaefer, Raymond F., Loftin, Laurence K., Jr., and Horton, Elmer A.:\nTwo-Dimensional Investigation of Five Related NACA Airfoil Sections\nDesigned for Rotating-Wing Aircraft. NACA TN 1922, 1949.\n\n4. Loftin, Laurence K., Jr., and Smith, Hamilton A.: Aerodynamic\nCharacteristics of 15 NACA Airfoil Sections at Seven Reynolds\nNumbers from $0.7 \\times 10^6$ to $9.0 \\times 10^6$. NACA TN 1945, 1949.\n\n5. Von Doenhoff, Albert E., and Abbott, Frank T., Jr.: The Langley\nTwo-Dimensional Low-Turbulence Pressure Tunnel. NACA TN 1283,\n1947.\n\n6. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.:\nSummary of Airfoil Data. NACA Rep. 824, 1945.", "timestamp": "2026-07-22T04:55:08.751601+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 32, "total_pages": 62, "image_filename": "19930082485_p32.jpg", "text": "NACA TN No. 1810\n\nSurface critical velocity ratio, $\\frac{V}{V_{cr}}$\n\nPressure side, $S_2$, in.\n\nSuction side, $S_1$, in.\n\nCalculated\nExperimental\n\nBlade-surface length, S, in.\n\n(b) Pitch section.\n\nFigure 2.- Continued. Blade-surface velocities at design conditions.\n\n[Figure: Graph showing surface critical velocity ratio vs. blade-surface length for pressure and suction sides, with calculated (circles) and experimental (squares) data points.]\n\nNACA\n\n31", "timestamp": "2026-07-22T04:55:08.960111+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 30, "total_pages": 33, "image_filename": "19930082487_p30.jpg", "text": "28\nNACA TN No. 1813\n\n| $\\alpha$ (deg) | Pressure distributions | Schlieren photographs |\n| :--- | :--- | :--- |\n| -2 | ——— | $\\circ$ |\n| 0 | - - - - - - | $\\triangle$ |\n| 2 | ——— | $\\square$ |\n| 4 | - - - - - - | $\\diamond$ |\n| 6 | ——— | $\\nabla$ |\n| 8 | - - - - - - | |\n\nUnpublished German pressure distributions; $\\alpha, 0^\\circ$.\n\nLocation of sonic point, $x/c$\nFree-stream Mach number, $M_o$\n\n$(x/c)_\\beta$\n$M_d$\n\nFigure 8.—Boundaries of supersonic region on upper surface of NACA 23015 airfoil section as a function of free-stream Mach number for various angles of attack.", "timestamp": "2026-07-22T04:55:12.208374+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 10, "total_pages": 66, "image_filename": "19930082914_p10.jpg", "text": "NACA TN No. 1857\n\nelectrodes are magnesium rods of $\\frac{1}{4}$-inch diameter with rounded tips approximately 1 inch apart. Each magnesium rod is held in a concentric hollow brass rod. A press fit is used for mechanical strength, and where the magnesium rod enters the brass rod, a weld is made. Each brass rod is soldered to a thin copper sheet, which extends for about an inch and is then soldered to a terminal on each of two condensers. Each of the two condensers, which are thus connected in parallel, is a 15,000-volt pyranol condenser and has, according to the manufacturer's specifications, a capacitance of 1.0 microfarad and an inductance of 0.5 microhenry. The condensers are charged, through a 10 megohm isolating resistor, to 16,000 volts. The discharge is initiated by means of a tickler spark between one of the magnesium electrodes and an auxiliary magnesium electrode placed between the other two and slightly to one side. The tickler spark is produced by the potential across the secondary of an automobile ignition coil that occurs when the 6-volt circuit through the primary is broken. This arrangement gives satisfactory control over the main discharge and facilitates synchronizing it with the opening of the camera shutter. The duration of the light from the main discharge was measured with a rotating-mirror apparatus and was found to be of the order of 3 microseconds.\n\nThe monochromator was constructed from a Bausch and Lomb \"Simplified Constant Deviation Prism Type\" Laboratory Spectrometer. The collimator objective, of aperture $f/8$, and the entrance slit of the original spectrometer were removed, and were replaced by a system of greater aperture that was constructed from available lenses. The lens $L_3$ (fig. 8) is an $f/1.6$ Kodak Anastigmat of 50-millimeter focal length. This lens focuses the light from the spark discharge on the adjustable entrance slit. (No pinhole is used at the spark.) The light is then made into a parallel beam by the lens $L_4$, which is an $f/2.3$ Bausch and Lomb Baltar of 2-inch focal length. The parallel beam then goes through the constant-deviation prism of the Pellin-Broca type. The beam is then focused on the adjustable exit slit by the lens $L_5$, which is an $f/8$ lens of 6-inch focal length and is the original lens furnished with the spectrometer. The exit slit is at the focus of the mirror $I_1$. The light from the slit is turned at right angles by a small prism, which is placed slightly off the axis of $I_1$. The mirror $I_1$ is a parabolic mirror of 4-inch diameter and 36-inch focal length and is front-surfaced with chrome-aluminum.\n\nIt is realized that considerably more light could be obtained from the system if the f-numbers of the various lenses and the mirror were properly matched, in order that the image of the entrance slit, which falls on the exit slit, would be more nearly the same size as the entrance and the exit slits. The system as used at present, however, works satisfactorily, and therefore no changes in it are contemplated.", "timestamp": "2026-07-22T04:55:12.409261+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 23, "total_pages": 37, "image_filename": "19930082450_p23.jpg", "text": "22\nNACA TN No. 1778\n\n$$ \\frac{P_l}{t_S}, \\text{ksi} $$\n\n$$ \\frac{H}{t_W} = 21 $$\n$$ \\left( \\frac{b_W}{t_W} = 20 \\right) $$\n\n$$ \\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S} $$\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n$$ 12.1 $$\n$$ 7.8 $$\n\nColors indicate minimum-weight proportions for $$ \\frac{t_W}{t_S} = 0.63 $$\nRed means some other blue means no other value of $$ \\frac{t_W}{t_S} $$ gives less weight\n\n31\n(30)\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n$$ 12.1 $$\n$$ 7.8 $$\n\n41\n(40)\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n$$ 12.1 $$\n$$ 7.8 $$\n\nNACA\n\n$$ \\frac{P_l}{L/t_S^2}, \\text{ksi} $$\n\nFigure 3: Direct-reading design chart for 24S-T aluminum-alloy Z-stiffened panels, $$ \\frac{t_W}{t_S} = 0.63 $$.", "timestamp": "2026-07-22T04:55:14.242435+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 2, "total_pages": 20, "image_filename": "19930085536_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:55:15.120730+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 6, "total_pages": 36, "image_filename": "19930085487_p6.jpg", "text": "position about one-third the distance from the root to the nearest node for the third bending-mode vibration photographed. This position is approximately the location of failure in the seventh and tenth stages when distances are proportioned according to total blade length. Consequently, in bending vibration only the first or second modes could cause the failures. In the torsional modes of vibration (figs. 3 and 5), the maximum-stress areas indicated by the node position are located at the leading and trailing edges near the blade root. Theoretically, the maximum stress for an airfoil-shaped bar is along the maximum blade thickness provided that the ends are permitted to warp at will. When the blade is rigidly restrained, however, the maximum-stress regions originate at the leading and trailing edges near the root and shift toward the center of the chord farther along the length of the blade. This conclusion was verified by data obtained from numerous resistance-wire strain gages cemented near a blade root in three directions of orientation, and by fatigue failing several blade specimens in torsion with a pneumatic vibrator. Torsional modes can therefore be disregarded because torsion would result in fatigue starting at the leading and trailing edges rather than at the maximum blade thickness. The node shapes on the blades of the seventh and ninth stages are distorted and the torsional modes therefore must be considered (fig. 8). The significance of the torsional modes might be more accurately evaluated from strain-gage data taken during engine operation.\n\nCritical-Speed Diagrams\n\nThe natural frequencies existing in each of the 10 stages of the compressor rotor, plotted on semilogarithmic grid to emphasize the more easily excited modes of vibration, are shown in the critical-speed diagrams of figure 9. The highest first bending-mode frequency corrected for the effect of centrifugal force is used for all diagrams. The exciting forces caused by wakes from the front-bearing-support arms and the stator blades and the effect of the split compressor case are shown plotted against rotor speed. The intersections of the curves of the natural frequencies and of those of the exciting-force frequencies indicate the critical speeds at which blade vibration resonance can be expected. Intersections below 14,000 rpm are ignored because extended operation of the engine is accomplished only at high speeds and the low speeds are quickly passed when starting the engine.\n\nA critical speed of 14,400 rpm is indicated on figure 9(a) at a third bending-mode vibration excited in the first-stage rotor blades by the 22 stator blades immediately following. High modes of vibration such as the third bending mode are difficult to excite and are therefore probably not serious. This critical speed is the only one readily apparent in the compressor.", "timestamp": "2026-07-22T04:55:22.896801+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 29, "total_pages": 41, "image_filename": "19930082476_p29.jpg", "text": "NACA TN No. 1801\n\n[Figure: Photograph of a model airplane mounted on a stand inside a tunnel, with a ruler for scale and a label reading \"NACA L-47679\" in the lower right corner.]\n\nFigure 2.- Photograph of the model as tested in the Langley 20-foot free-spinning tunnel.\n\n27", "timestamp": "2026-07-22T04:55:23.930247+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 40, "total_pages": 66, "image_filename": "19930082245_p40.jpg", "text": ".16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.56\n-.60\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section hinge-moment coefficient, $c_h$\nNACA\nNACA TN No. 1596\n39\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n$\\delta_a$\n(deg)\n30\n18\n12\n4\n2\n0\n-4\n-6\n-12\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nAileron section normal-force coefficient, $c_{n_a}$\n\n(g) $c_n=0.7$.\nFigure 7.-Continued.", "timestamp": "2026-07-22T04:55:27.017758+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 16, "total_pages": 53, "image_filename": "19930082542_p16.jpg", "text": "NACA TN No. 1867\n15\n\nof the treatments on bar stock. This table shows that the discs tend to have lower properties than bar stock; but the actual results from bar stock and from estimates based on hardness are of the same order as the test results from the discs.\n\nThe trends found for the types of treatment should be applicable to other alloys of the same general type. In each case, however, additional test work will have to be carried out in order to establish the optimum treatments for other alloys. Relative stability of the structures produced will probably vary for each analysis. The types of precipitates which form may require other conditions of heat treatment and hot-cold-work for best properties. The data for low-carbon N-155 alloy in this report should serve as a guide to reduce the testing of other alloys to a minimum.\n\nTheoretical Considerations\n\nThe mechanism by which the treatments influence properties has not yet been established. Pronounced changes in precipitated constituents occur during solution and aging treatments. The most logical explanation is that the treatments change the size, dispersion, and possibly the composition of the precipitates which form during the treatments or during testing. Strain hardening may also be an important part of the pronounced effect of hot-cold-work.\n\nThe similarity in properties resulting from cold-working at room temperature and hot-cold-working at $1200^\\circ$ F suggests that precipitation during hot-cold-work at temperatures up to $1200^\\circ$ F is not a major factor. If this is the case then major effects of hot-cold-work must be through strain hardening and its influence on precipitation reactions during testing. If precipitation during hot-cold-working is appreciable then a similar amount of precipitation must occur very rapidly during testing after cold-working at room temperature. The pronounced influence of prior treatment on the properties after hot-cold-working would seem to indicate that precipitation reactions are important. In view of the similarity of the slopes of the curves of stress against rupture time after hot-cold-working and after solution-treating it seems unlikely that strain hardening alone could be the predominating factor.\n\nThe data are not complete enough to indicate how much precipitation during hot-cold-working at higher temperature influenced the properties. The indications are, however, that above $1400^\\circ$ F aging occurs which is not so beneficial to strength as that which occurs after working at lower temperatures.\n\nThe relatively steep slope of the curves of stress against rupture time after most aging treatments when compared with those for solution-treated materials suggests that the precipitates which form during aging are either unstable or are distributed or formed in a manner which", "timestamp": "2026-07-22T04:55:28.690162+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 23, "total_pages": 50, "image_filename": "19930082496_p23.jpg", "text": "22\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T04:55:29.524064+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 14, "total_pages": 78, "image_filename": "19930082618_p14.jpg", "text": "12\nNACA TN 1945\n\ncomparable measurements of the lift-curve slope for different Reynolds numbers did not appear feasible by the method employed in reference 1. The lift-curve slopes were therefore considered to be defined by the best straight line through the experimental points between zero lift and the design lift coefficient for the cambered airfoils. For the two symmetrical sections, the lift-curve slopes were determined by the best fairing of the data from zero lift to a few tenths in lift coefficient above and below zero lift. The lift-curve slopes corresponding to all the Reynolds numbers from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ were measured according to this procedure and are presented for the 15 airfoils in the smooth and rough surface conditions in figure 17.\n\nAn examination of the data of figure 17 indicates that the value of the slope of the lift curve for the smooth airfoils decreases as the Reynolds number is lowered from $9.0 \\times 10^6$ to $0.7 \\times 10^6$. The magnitude and character of the scale effect vary somewhat for the different airfoils; however, these variations in scale effect do not form any consistent trends with systematic changes in the design parameters of the airfoils. In most instances, for the smooth airfoils, the decrease in lift-curve slope which accompanies reductions in Reynolds number is greatest between Reynolds numbers of $3.0 \\times 10^6$ and $0.7 \\times 10^6$, with variations in Reynolds number from $9.0 \\times 10^6$ to $3.0 \\times 10^6$ usually having an almost imperceptible effect on the slope of the lift curve (fig. 17(a)). In comparison with the data for the NACA 6-series airfoils, the lift-curve slope of the NACA $64_1A212$ airfoil is seen to be rather low at all Reynolds numbers. As pointed out in reference 3, the trailing-edge angles of the NACA 6A-series sections, which are larger than the trailing-edge angles of the NACA 6-series sections, cause reductions in the lift-curve slope.\n\nThe addition of roughness to the leading edge usually results in lower lift-curve slopes for all the airfoils (fig. 17(b)). In general, however, for any particular airfoil, the decrement in lift-curve slope due to roughness does not seem to vary to any large extent with the Reynolds number.\n\nAngle of zero lift.- The data presented in figure 18 indicate that the angle of zero lift of most of the airfoils changes to some small extent with Reynolds number, but in most cases the scale effect is relatively insignificant. The addition of standard leading-edge roughness (fig. 18(b)) causes a change in the magnitude of the angle of zero lift of most of the airfoils.\n\nMaximum lift coefficient.- The lift parameter which is most affected by variations in the Reynolds number is the maximum lift coefficient (figs. 1 to 15). A discussion of the flow phenomena associated with the occurrence of maximum lift and the relationship between these phenomena and the Reynolds number is given in reference 5.", "timestamp": "2026-07-22T04:55:32.181249+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 5, "total_pages": 46, "image_filename": "19930085519_p5.jpg", "text": "4\nNACA RM No. L8K19\n\nc\nlocal wing chord measured along lines parallel to X-axis, feet\n\ny\nlateral distance from plane of symmetry along Y-axis, feet\n\nV\nfree-stream velocity, feet per second\n\n$V_s$\nsinking velocity, feet per second\n\n$V_g$\ngliding velocity, miles per hour\n\n$\\rho$\nmass density of air, slugs per cubic foot\n\n$\\alpha$\nangle of attack with respect to chord plane of model, degrees\n\n$\\delta_p$\nplug-aileron projection, percent local wing chord, negative\nwhen plug is projected above wing upper surface\n\n$\\delta_a$\naileron deflection measured in planes perpendicular to aileron\nhinge axis, degrees\n\n$\\delta_f$\nflap deflection measured in planes perpendicular to flap leading\nedge, degrees\n\nR\nReynolds number\n\nThe rolling-moment and yawing-moment coefficients represent the\naerodynamic effects that occur on a complete wing as a result of deflection\nof the control on one semispan of the complete wing; the lift, drag, and\npitching-moment coefficients represent the aerodynamic effects that occur\non the complete wing as a result of deflection of the lift flap on both\nsemispans of the complete wing.\n\nThe test data have been corrected for blockage and jet-boundary\neffects, including the reflection-plane corrections to the rolling-moment\nand yawing-moment coefficients. The variation of the corrections to the\nrolling-moment and yawing-moment coefficients with span of the lateral-\ncontrol device is presented in reference 1. The rolling-moment and yawing-\nmoment coefficient corrections applied to the data presented herein were\ntaken directly from reference 1 for the span of the control device under\nconsideration.\n\nNo corrections were made to the data to account for wing twist caused\nby control deflections or projections or flap deflection.", "timestamp": "2026-07-22T04:55:38.110678+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 37, "total_pages": 47, "image_filename": "19930093773_p37.jpg", "text": "36\nNACA RM E9G09\n\nCorrected air flow, $W_a\\sqrt{\\theta}/\\delta$, lb/sec\n\n| Altitude (ft) | |\n| :--- | :--- |\n| $\\circ$ | 5,000 |\n| $\\square$ | 15,000 |\n| $\\diamond$ | 25,000 |\n| $\\triangle$ | 35,000 |\n| $\\nabla$ | 45,000 |\n| $\\triangledown$ | 50,000 |\n\n[Figure: Graph plotting Corrected air flow against Corrected engine speed with multiple curves representing different altitudes]\n\nCorrected engine speed, $N/\\sqrt{\\theta}$, rpm\n\n(b) Air flow.\n\nFigure 6. - Continued. Effect of altitude on variation of corrected engine performance with corrected engine speed at flight Mach number of 0.21.\n\n1159", "timestamp": "2026-07-22T04:55:38.999242+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 16, "total_pages": 30, "image_filename": "19930082585_p16.jpg", "text": "NACA TN 1907\n15\n\nThe various ways in which the pitch is changed, for this investigation, are illustrated in figure 2. The different curves correspond to different values of the constants $A_1$, $A_2$, $a_1$, and $a_2$ (equation (2)).\n\nThe variations of induced velocity with time, given by equation (1), are illustrated in figure 3, for $k = 1$, and, in the cases of pitch reduction after power failure, for $k = \\infty$.\n\nRESULTS AND DISCUSSION\n\nProbable Effect of Assuming the Variation of\n\nInduced Velocity with Time\n\nA typical variation of descending velocity $V$ and rotor angular velocity $\\Omega$ with time after power failure is illustrated in figure 4. These calculations were made by the step-by-step process given in the section entitled \"Solution for $\\Omega$ and $V$,\" neglecting flapping. Two calculations were made, for $k = 1.0$ and $\\infty$, in order to evaluate the importance of $k$.\n\nSince, in the usual case, the initial and final values for the induced velocity are not very different, it seems intuitively evident that the exact variation of $v(t)$ ought not to affect the variations of $V(t)$ and $\\Omega(t)$ appreciably. This is confirmed by figure 4, where $V(t)$ and $\\Omega(t)$ are compared for the widely different values of $k$, representing the variations of $v(t)$ shown in figure 3.\n\nIt may be observed that, if the induced velocity varied in such a manner that its value was not always between the initial and final values, then it might be important to consider the actual variation. In that case, $v(t)$ would have to be determined from experiment. It should be remarked that, in the solution for $\\Omega(t)$ and $V(t)$ just referred to, any suitable variation of $v(t)$ may be assumed. The exponential variation was convenient in the analytical solution for $\\beta(t)$.\n\nInterpretation of the Different Rates of Pitch\n\nChange Considered\n\nThe extremes of the types of pitch change considered are the uppermost and lowermost curves of figure 2, representing no pitch change and instantaneous pitch reduction at the instant of power failure. The case of no pitch change could occur with rigid blades (or hinged blades so", "timestamp": "2026-07-22T04:55:41.761679+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 8, "total_pages": 28, "image_filename": "19930085471_p8.jpg", "text": "6\nUNCLASSIFIED\nCONFIDENTIAL\nNACA RM No. L8J11\n\nobserved in the tests that the thick airfoils tended to diverge even though the elastic axis was ahead of the 50-percent-chord position.\n\nSince practical winged vehicles pass through the subsonic speed range in order to reach supersonic speeds, some discussion of and comparison with subsonic criterions are desirable. For this purpose, incompressible flutter-speed coefficients were computed by the method of reference 2 for which first bending and uncoupled first torsion frequencies and damping coefficients $\\xi_h$ and $\\xi_\\alpha$ were utilized. That flutter-speed coefficients based on subsonic theory are conservative with respect to supersonic speeds has been suggested in reference 3 and, in fact, is indicated by some of the calculations in reference 1. This is also indicated by the present tests, as illustrated in figure 8 in which the experimental flutter-speed coefficients are plotted against the incompressible theoretical flutter-speed coefficients. The statement may not be true in general; for example, the condition when the frequency ratio $\\frac{\\omega_h}{\\omega_\\alpha} = 1$ may need further investigation and, in any case, the margin of safety is not established. Some of the models were permitted to encounter the tunnel transient speeds and, for example, model \"F\" which had fluttered at Mach number 1.3 was held in the tunnel while the tunnel was brought up to speed. The wing fluttered and broke at a Mach number of about 0.7, a result which is in general agreement with the subsonic calculation. Flutter data obtained with bombs and rocket missiles (references 4 to 6) and other experience indicate that if flutter failures occur, they occur, in general, at near sonic speeds. For the practical purpose of making preliminary estimates of a wing flutter speed such formulas as, for example, the approximate flutter formula in reference 2, or similar criterions, thus appear useful over a wide range of speeds.\n\nIn reference 3, Smilg suggests a torsional frequency criterion $\\omega_\\alpha c_w > 1000$ feet per second as sufficient to prevent one degree of torsional flutter at supersonic speeds. The criterion is based on the assumption that the first bending frequency is very high with respect to the first torsion frequency. In order to look into this criterion a cantilever model was equipped with tip weights at both the leading and trailing edges which reduced the torsional frequency. The results of the tests are shown in table II. In all cases a slight shift of the center of gravity ahead of the location at flutter stopped the flutter. The fact that flutter is extremely sensitive to the center-of-gravity position and that the values of the product $\\omega_\\alpha c_w$ are far below 1000 indicates that the criterion is overly conservative when applied to cantilever wings with normal bending-torsion frequency ratios. The data further suggest that for cantilever wings the bending degree of freedom may suppress the one-degree-of-freedom torsional flutter and that bending-torsion effects occur.\n\nCONFIDENTIAL\nUNCLASSIFIED", "timestamp": "2026-07-22T04:55:44.237445+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 25, "total_pages": 99, "image_filename": "19930082511_p25.jpg", "text": "NACA TN No. 1826\n\n$$\n1 - \\frac{-(A + Bz_1)\\sqrt{1 - z_1^2} - \\frac{2(A + Bz_1)z_1^2}{\\sqrt{1 - z_1^2}} + 2Bz_1\\sqrt{1 - z_1^2}}{2z_1}\n$$\n\nThis fraction can be greatly simplified by use of equation (4), and the result, added to the previously derived limit, gives the desired correction at the vortex:\n\n$$\n\\lim_{\\xi \\to \\xi_1} q_1(\\xi, \\xi_1) = q_1(\\xi_1) = i \\left[ -\\frac{\\Gamma'}{4} + \\frac{\\Gamma'}{2(1 - z_1^2)} + B\\sqrt{1 - z_1^2} + \\frac{\\Gamma' z_1}{4iy_1} \\right] \\tag{8}\n$$\n\nFor the special case in which the vortex is on the tunnel axis ($z = iy$), this expression reduces to the following form (after substituting for $B$ from equation (6)):\n\n$$\nq_1\\left(\\xi + \\frac{1}{2}\\right) = -\\frac{i\\Gamma'}{2} \\frac{y_1^2}{1 + y_1^2} \\tag{9}\n$$\n\n**Upstream perturbation velocity.** — If the vortex is not on the tunnel axis, $A$ will not be zero (compare equation (6)). The tunnel interference velocity far upstream in the closed part of the tunnel is found by putting $\\xi = -\\infty$ and $z = 0$ in equation (5), which then reduces to\n\n$$\nq_1(-\\infty, \\xi_1) = iA\\left(\\frac{1}{\\bar{z}_1} - \\frac{1}{z_1}\\right) = -\\frac{2Ay_1}{|z_1|^2}\n$$\n\nwhich is real. For this unsymmetrical case, therefore, a finite longitudinal perturbation velocity is found far upstream in the closed part. As was pointed out in part I, such results appear because the problem was set up so that the longitudinal perturbation velocity on the open boundary is zero. If the velocity far upstream in the closed part is to be taken as the base, the result means merely that the longitudinal velocity on the open boundary exceeds this base velocity by $\\frac{2Ay_1}{|z_1|^2}$ and that there is a corresponding difference in pressure between the closed part and the space surrounding the jet.", "timestamp": "2026-07-22T04:55:47.427269+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 16, "total_pages": 58, "image_filename": "19930082617_p16.jpg", "text": "```markdown\nNACA TN 1962\n15\n\n| Stringers | |\n| :--- | :--- |\n| O | 1 to 9 |\n| X | 10 to 16 |\n\n| Moment (in. - lb) | |\n| :--- | :--- |\n| 1 | 28.8 X $10^3$ |\n| 2 | 57.6 X $10^3$ |\n| 3 | 86.4 X $10^3$ |\n| 4 | 115.2 X $10^3$ |\n| 5 | 144.0 X $10^3$ |\n| 6 | 201.6 X $10^3$ |\n\n3.86\"\n[Figure: Cross-section diagram showing Band N and A-A]\n\n| Distance from horizontal diameter, in. | Strain |\n| :--- | :--- |\n| 10 | |\n| 9 | |\n| 8 | |\n| 7 | |\n| 6 | |\n| 5 | |\n| 4 | |\n| 3 | |\n| 2 | |\n| 1 | |\n| 0 | 20 16 12 8 4 -4 -8 -12 -16 -20 X $10^{-4}$ |\n| 1 | |\n| 2 | |\n| 3 | |\n| 4 | |\n| 5 | |\n| 6 | |\n| 7 | |\n| 8 | |\n| 9 | |\n| 10 | |\n\n[Figure: Strain diagram curves labeled 1 through 6]\n\nNACA\n\nFigure 4.- Strain diagram of cylinder 72. Band N.\n```", "timestamp": "2026-07-22T04:55:48.381742+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 31, "total_pages": 33, "image_filename": "19930082487_p31.jpg", "text": "NACA TN No. 1813\n29\n\n.3\nNACA 4415\n.2\nMd\n.1\n.05\n.6 .7 .8\nMach number, M₀\n\nNACA 0015\nMd\n.6 .7 .8\n\nLocation of forward sonic point, x/c\n.2\nNACA 23015\nMd\n.1\n.04\n.5 .6 .7 .8\nMach number, M₀\n\nα (deg)\n○ -2\n+ 0\nx 2\n□ 4\n◇ 6\n△ 8\n\n.4\nNACA 65₂-215, a=0.6\n.3\nMd\n.2\n.1\n.06\n.7 .8\nMach number, M₀\n\nNACA 66,2-215, a=0.6\nMd\n.6 .7 .8\nMach number, M₀\n\nNACA\n\nFigure 9.— Location of forward sonic point as a function of free-stream Mach number for several NACA airfoil sections.", "timestamp": "2026-07-22T04:55:51.255009+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 15, "total_pages": 36, "image_filename": "19930082614_p15.jpg", "text": "NACA TN 1939\n\n$\\frac{1}{57.3} \\frac{\\mathrm{d} \\gamma}{\\mathrm{~d} t}=\\frac{\\mathrm{V}}{\\mathrm{r}}$\n\nSubstituting $\\frac{\\mathrm{V}}{\\mathrm{r}}$ from equation (9),\n\n$\\frac{\\mathrm{d} \\gamma}{\\mathrm{dt}}=57.3 \\mathrm{~g} \\frac{\\mathrm{n}-\\cos \\gamma}{\\mathrm{V}}$\n\nfrom which, by using small finite time intervals, $\\Delta \\gamma$ may be computed as\n\n$\\Delta \\gamma=\\frac{\\mathrm{d} \\gamma}{\\mathrm{dt}} \\Delta \\mathrm{t}$\n\nThe estimated average speed $\\bar{V}_{e}$ during the time $\\Delta t$ is used in calculations for radius of curvature and $d \\gamma / d t$. In addition, an estimate of the average flight-path inclination is necessary to determine the value of the abscissa $(n-\\cos \\bar{\\gamma})$. The accuracy of this estimate can be checked as soon as $\\Delta \\gamma$ has been found, if it is assumed that $\\bar{\\gamma}$ is the initial angle plus half the increment.\n\nFigure 5(f) shows the variation with $\\Delta \\gamma$ of the quantity $\\left[\\cos \\gamma_{1}-\\cos \\left(\\gamma_{1}+\\Delta \\gamma\\right)\\right]$. There are two sets of values of $\\gamma_{1}$ identifying the curves, one of which applies when $\\Delta \\gamma$ is positive and one when $\\Delta \\gamma$ is negative. This dual labeling is used because the curves have been plotted on only one side of the vertical axis. The algebraic sign of $\\Delta \\gamma$ may readily be ascertained since it is the same as the sign of the quantity $(n-\\cos \\gamma)$. When the ordinate of figure 5(f) and the radius of flight-path curvature, figure 5(b), are known, the change in altitude is given by figure 5(e). Although the algebraic sign of the change in altitude is not shown in the graph, in most cases it is evident from the problem. If the sign is not apparent, a simple diagram showing $\\gamma_{1}$ and $\\Delta \\gamma$ will indicate whether the altitude increases or decreases during the time $\\Delta t$.\n\nIt is seen that figure 5 cannot be used when the flight-path radius becomes very large. In this case the altitude change may be calculated with sufficient accuracy from equation (7).\n\nThe graphical solution for the change in altitude is arranged so as to be used directly with maneuvers in a vertical plane which are identified by the magnitude of the normal acceleration or load factor. For maneuvers not wholly in a vertical plane, part of the total load factor results from accelerations in a horizontal direction and does not affect the altitude. In this case the value of the load factor to", "timestamp": "2026-07-22T04:55:59.012450+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 11, "total_pages": 66, "image_filename": "19930082914_p11.jpg", "text": "10\nNACA TN No. 1857\n\nThe light from the green triplet of magnesium is used, which has wavelengths of 5167, 5173, and 5184 angstrom units. Of course, at the high exciting voltage that is used, the light does not consist only of these three wavelengths but is nearly a continuous spectrum. It was found that the light from this region produced much more satisfactory interference fringes than that from any other region of the spectrum of a high-voltage magnesium spark. (The Princeton group reported (reference 10) that they obtained best results by using the blue line at 4481 angstroms.) The monochromator was set at 5170 angstroms, and the slits were set at a width of 0.3 millimeter. A band about 30 angstroms wide was passed by the exit slit, and about 180 usable fringes were obtained. It was found that the appearance, or contrast, of the fringes could be improved by reducing the length of the exit slit to 3/32 inch.\n\nThe fringes were photographed with an Eastman Anastigmat aerial camera lens of 13.5-inch focal length and f/3.5 aperture. Kodak Linagraph Ortho film was used. The Linagraph Ortho is a very fast orthochromatic film of moderate contrast and high resolving power and is designed for photographing high-speed transient phenomena on green-fluorescing cathode-ray screens. Negatives of about $1\\frac{1}{2}$-inch diameter were taken. (Inasmuch as the film was used in the 35-mm size, a portion of the light did not hit the film. Since the interferograms shown in the present paper were taken, the film has been changed to the 70-mm size.) The moderate grain size of the film permitted enlargements of sixteen or more times the diameter of the negative, or seven or more times actual size.\n\nA Kodak Supermatic shutter was placed at the focus of the camera lens. The shutter was set at 1/50 second and was synchronized with the light-source spark. For convenience in moving the film, a slightly modified Argus C-3 camera, with the lens removed, was used as a film holder.\n\nThe procedure that was followed for each interferogram of the flow was to take first an interferogram with no flow, then one with a transparent ruler in the test section for establishing the scale, then one with flow, then a final one without flow. The film was developed in Kodak D-72 for the exceptionally long time of 14 minutes at 68° F.\n\nAdjustment of the interferometer.- When the interferometer is first set up, a number of adjustments must be made in order to obtain fringes and to orient them properly. Most of the adjustment procedure that was used is a more-or-less standard procedure, but part is original.\n\nThe first step in the initial adjustment of the interferometer is to make the reflecting surfaces of the two splitter plates and the two mirrors nearly parallel. This is done by making the two mirrors as", "timestamp": "2026-07-22T04:56:03.156553+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 30, "total_pages": 41, "image_filename": "19930082476_p30.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:56:04.054787+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 3, "total_pages": 20, "image_filename": "19930085536_p3.jpg", "text": "```markdown\nNACA RM No. E5K05\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nPRESSURE DISTRIBUTIONS ON THIN CONICAL BODY OF ELLIPTIC\nCROSS SECTION AT MACH NUMBER 1.89\n\nBy Stephen H. Maslen\n\nSUMMARY\n\nAn investigation was conducted to determine the pressure\ndistribution on a conical body of elliptic cross section at a\nMach number of 1.89. Experimental data are presented for a range\nof angles of yaw from -16° to 16° and angles of attack from -10°\nto 10°.\n\nAs the angle of flow deflection was increased, the deviation\nfrom experiment of the theoretical pressure distribution slightly\nincreased, although agreement was satisfactory over the entire\nrange of calculations. Comparison of the complete equation for\npressure coefficient (that is, the equation including all the per-\nturbation velocity components) with the equation usually used in con-\nnection with the linearized theory indicated that the terms usually\nneglected appreciably alter the predicted values of the pressure\ncoefficient. Although the complete equation gave better agreement\nwith experiment for the elliptic cone investigated than did the\nlinearized equation, the opposite result was found when a similar\ncomparison with the exact results of Taylor and Maccoll was made.\nThe excellent agreement between experiment and linearized theory\nmay therefore be fortuitous.\n\nINTRODUCTION\n\nAircraft designers are currently in need of a reliable means\nof estimating loads on body contours that might be used as fuse-\nlages of supersonic airplanes. Several methods have been available\nfor the theoretical calculation of force distribution over bodies\nof revolution, as well as considerable experimental data for check-\ning such calculations (for example, references 1 to 5). Recently,\na theoretical method for calculating the pressure distribution\nover conical bodies of noncircular cross section has also become\navailable (reference 6).\n```", "timestamp": "2026-07-22T04:56:04.595833+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 24, "total_pages": 50, "image_filename": "19930082496_p24.jpg", "text": "NACA TN No. 1836\n23\n\n[Figure: A tensile specimen with labels pointing to \"Asbestos wrapping\" and \"Strain gages\". A scale marked \"INCHES\" is visible. In the bottom right corner, there is a NACA logo with the text \"C-21896\" and \"7-27-48\".]\n\nFigure 2. - Tensile specimen before investigation.", "timestamp": "2026-07-22T04:56:11.257375+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 19, "total_pages": 49, "image_filename": "19930082498_p19.jpg", "text": "```markdown\n18\n\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n<!-- Table (69, 199, 888, 817) -->\n\\begin{tabular}{|c|l|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c", "timestamp": "2026-07-22T04:56:13.374128+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 17, "total_pages": 53, "image_filename": "19930082542_p17.jpg", "text": "16\nNACA TN No. 1867\n\ndoes not provide as good strength as those which precipitate during testing. The aging treatments which were effective in changing the properties were at considerably higher temperatures than the test temperature of 1200° F. It therefore appears that the higher temperature treatments overage the alloy in comparison with the aging which takes place during prolonged rupture testing. The high strength at long time periods predicted by the relatively flat curves of stress against rupture time after solution treatment indicates that the precipitates formed during testing are distributed or formed in a manner more advantageous to strength, or of a different, more effective type, than those formed during aging treatments. The very low rupture strength at short time periods after solution treatments suggests that precipitation takes place very slowly at 1200° F and that considerable time must elapse before sufficient precipitation occurs to develop strengths equivalent to those obtained by other added treatments.\n\nThe low ductility in the rupture test and sensitivity to brittleness at stress concentrations associated with the completely solution-treated condition are also probably manifestations of precipitation during testing. The improvement in ductility and toughness through subsequent aging treatments probably results from these treatments changing the precipitate size and dispersion.\n\nIt is difficult to understand why aging treatments alone do not have more influence on the properties at room temperature. Apparently the precipitates are of a form or dispersion which do not have much effect on properties at room temperature. An unlikely alternative explanation could be that little actual precipitation occurs during the heat treatments and that the major effect of the treatments is observed at high temperatures because of their influence on precipitation during testing.\n\nIn summary then, the probable mechanism which controls rupture properties involves the following conceptions: (1) Solution treatments remove strengthening due to precipitation or strain hardening. Consequently, properties are low at room temperature and for short time periods at 1200° F. During prolonged testing at 1200° F, however, precipitation occurs slowly and develops high strength. (2) Aging treatments cause either a different form or type of precipitate to appear which gives higher short-time rupture strengths than the plain solution-treated material. The longer-time strength, however, obtained with the initial precipitation of aging is not so high as that which results from precipitation during testing of the material solution-treated only. The precipitation during aging has relatively little effect on properties at room temperature possibly because the aging treatment mainly influences precipitation during testing at high temperatures. (3) Hot-cold-work provides high strength at short time periods at 1200° F and at room temperature possibly from strain hardening. High strength in prolonged rupture tests under proper conditions of heat treatment and", "timestamp": "2026-07-22T04:56:17.977005+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 7, "total_pages": 36, "image_filename": "19930085487_p7.jpg", "text": "NACA RM No. E9J22\n\nWithin the operating range of the engine, the exciting force required to cause a resonant vibration in the seventh-stage blades in the first bending mode is approximately the fourth order of the rotor speed (fig. 9(g)). The only fourth-order exciting force obviously present in the compressor is caused by the four arms supporting the front main bearing but it is doubtful that the effect of the arms would carry through as far as the seventh stage. In the tenth stage, the resonant point of first bending-mode vibrations at 16,600 rpm coincides exactly with a sixth-order excitation (fig. 9(j)). The blades are possibly being mechanically or aerodynamically excited but no exciting force of these orders could be determined. All the failures reported occurred in complete engines and, although a compressor has been operated at various speeds up to 17,000 rpm in an NACA test cell for more than 450 hours, no failures have occurred. Some coupling effect might possibly take place between the turbine or the accessories and the compressor during complete engine operation. Operation at high pressure ratios may also produce the air-flow velocities and pressure conditions necessary to excite blade flutter.\n\nSUMMARY OF RESULTS\n\nResults of an investigation to determine the reason for failure of blades of the seventh and tenth compressor stages in the 10-stage axial-flow compressor at a speed of 16,600 rpm indicated a possibility of exciting first bending-mode vibrations in the seventh-stage blades by a fourth-order excitation of the rotative speed. The tenth-stage blades were possibly excited in the first bending mode by a sixth-order excitation.\n\nAnother possible source of excitation was the existence of an aerodynamic condition causing a flutter form of blade vibration.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.", "timestamp": "2026-07-22T04:56:20.736552+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 38, "total_pages": 47, "image_filename": "19930093773_p38.jpg", "text": "NACA RM E9G09\n37\n\nCorrected fuel consumption, $W_f/\\delta\\sqrt{\\theta}$, lb/hr\n\n| Altitude (ft) | |\n| :--- | :--- |\n| $\\circ$ | 5,000 |\n| $\\square$ | 15,000 |\n| $\\diamond$ | 25,000 |\n| $\\triangle$ | 35,000 |\n| $\\nabla$ | 45,000 |\n| $\\triangleright$ | 50,000 |\n\n[Graph showing Corrected fuel consumption vs Corrected engine speed with multiple curves for different altitudes]\n\nCorrected engine speed, $N/\\sqrt{\\theta}$, rpm\n\n(c) Fuel flow.\nFigure 6. - Continued. Effect of altitude on variation of corrected engine performance with corrected engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:56:23.400321+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 15, "total_pages": 78, "image_filename": "19930082618_p15.jpg", "text": "NACA TN 1945\n\nThe plot of maximum lift against Reynolds number for the different airfoils (figs. 19 to 22) shows that, in all cases, decreasing the Reynolds number from the highest value to $0.7 \\times 10^6$ effects reductions in the maximum lift of the airfoils, with and without split flaps, in both the smooth and rough surface conditions. The manner in which the maximum lift of the airfoils varies with Reynolds number and the magnitude of this variation are seen to depend upon the airfoil design, surface condition, and whether a split flap is employed. Unfortunately, the data also show that the type and magnitude of the scale effect on the maximum lift do not vary in any very consistent manner with the airfoil-design parameters investigated. It is not possible, therefore, to predict from the comparative values of the maximum lift of a group of airfoils at one Reynolds number the advantage one airfoil will have over another at any other Reynolds number.\n\nAs an example of the manner in which the merits of different airfoils change with Reynolds number consider the manner in which the comparative values of the maximum lift of the NACA 64-409 and NACA 643-418 airfoils in the smooth condition change as the Reynolds number is lowered from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ (fig. 19). Notice also that the rather large advantage of the NACA 23012 airfoil in the smooth, plain condition as compared with the NACA 641-412 and NACA 4412 sections decreases and finally vanishes as the Reynolds number is progressively reduced from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ (figs. 20 and 22). In general, there is less scale effect on the maximum lift of the airfoils with rough leading edges than on the airfoils with smooth surfaces. Surface roughness, nevertheless, has a large effect upon the comparison of some of the airfoils for the different Reynolds numbers. Again consider the data for the NACA 23012 section which show that the maximum lift of this plain airfoil with roughness becomes progressively less favorable relative to that of comparable NACA 6-series sections (NACA 641A212 and NACA 641-412) as the Reynolds number is reduced and is actually less than that of the NACA 64-409 section below $2.0 \\times 10^6$; whereas at practically all Reynolds numbers, the plain NACA 23012 section with smooth surface has maximum lift coefficients as high as or higher than those of comparable 6-series airfoils. With split flaps deflected 60°, the data show that the amount and type of maximum-lift variation with Reynolds number are not necessarily the same as indicated by the results for the plain airfoils; and again, the comparative values of the maximum lift of the various airfoils with split flaps are seen to change with the Reynolds number and surface condition. From the viewpoint of the aircraft designer, the most important conclusion to be drawn from these maximum-lift data is that the selection of an airfoil for a given application must be made from data at a Reynolds number corresponding to the Reynolds number of the application.", "timestamp": "2026-07-22T04:56:24.131478+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 24, "total_pages": 37, "image_filename": "19930082450_p24.jpg", "text": "NACA TN No. 1778\n23\n\n$$\n\\frac{P_l}{l_S}, \\text{ksi}\n$$\n\n$$\n\\frac{H}{l_W} = .26\n$$\n$$\n\\left(\\frac{b_W}{l_W} = .25\\right)\n$$\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n12.1\n7.8\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n16.8\n12.1\n7.8\n\n$$\n\\sigma_{cr}, \\text{ksi}\n$$\n19.5\n16.4\n\n$$\n\\sigma_{cy} = 44 \\text{ ksi}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{e}}, \\text{ksi}\n$$\n$$\n\\frac{l_W}{l_S} = 0.63\n$$\n\nFigure 3-Concluded.\n\n[Figure: Three stacked charts plotting $\\frac{P_l}{l_S}$ vs $\\frac{P_l}{L\\sqrt{e}}$ with various curves and annotations. A cross-section diagram of a structural member is shown in the top chart.]", "timestamp": "2026-07-22T04:56:26.609175+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 33, "total_pages": 62, "image_filename": "19930082485_p33.jpg", "text": "32\n\nSurface critical velocity ratio, $\\frac{V}{V_{cr}}$\n\nPressure side, $S_2$, in.\n\nSuction side, $S_1$, in.\n\nCalculated\nExperimental\n\nNACA\n\nBlade-surface length, S, in.\n(c) Tip section.\nFigure 2.- Concluded. Blade-surface velocities at design conditions.\n\nNACA TN No. 1810", "timestamp": "2026-07-22T04:56:30.440120+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 6, "total_pages": 46, "image_filename": "19930085519_p6.jpg", "text": "NACA RM No. L5K19\n\nAPPARATUS AND MODEL\n\nThe right semispan sweptback-wing model was mounted in the Langley 300 MPH 7- by 10-foot tunnel as shown in figure 2. The root chord of the model was adjacent to the ceiling of the tunnel, the ceiling thereby serving as a reflection plane. The model was mounted on the balance system in such a manner that all forces and moments acting on the model could be measured. A small clearance was maintained between the model and the tunnel ceiling so that no part of the model came in contact with the tunnel structure. A root-fairing strip was attached to the model to deflect the air that flows into the tunnel test section through the clearance hole between the model and the tunnel ceiling so as to minimize the effects of any such inflow on the flow over the model.\n\nThe model had $42^\\circ$ of sweepback referred to the wing leading edge, an aspect ratio of 4.01, and was constructed of laminated mahogany to the plan form shown in figure 1. The airfoil section normal to the 0.272 chord line was constant throughout the span and was of NACA 64$_1$-112 airfoil profile. The tip of the wing was rounded off beginning at 0.975$\\frac{b}{2}$ in both plan form and cross section. The model had no geometric twist or dihedral.\n\nThe full-span 20-percent-chord slotted flap was built to the plan form and section dimensions shown in figures 3 and 4, respectively. The flap was fitted with an attachment bracket at three spanwise locations and each bracket could be adjusted to give several flap deflections and a range of positions of the flap nose with relation to the wing trailing-edge upper-surface lip. A partial-span slotted flap was formed by cutting the flap at the 51-percent-span station on a line parallel to the model plane of symmetry. The details of flap-slot flow-control vanes A and B investigated on the semispan-wing model are presented in figures 5 and 6, respectively. A half-span Zap-type flap investigated on the semispan-wing model was built of thin plywood and was deflected down $60^\\circ$ about a hinge line on the wing trailing edge (fig. 7). The small slot between the flap leading edge and the wing trailing edge was sealed.\n\nThe plan form and section dimensions of the basic plug ailerons investigated are shown in figures 3 and 8, respectively. The plug ailerons were built in six segments of $\\frac{1}{4}$-inch aluminum plate and had $\\frac{1}{8}$-inch-thick steel actuating arms screwed to the ends of each plug segment. A clamp was provided on each actuating arm to hold the plug aileron at the desired projection. The plugs could be adjusted through a range of projections from 0 percent to -7 percent of the local wing chord. In addition, the plug aileron was investigated with the slot lower lip refaired from the original sharp lip to a smooth air inlet as shown in figure 9.", "timestamp": "2026-07-22T04:56:34.795274+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 17, "total_pages": 30, "image_filename": "19930082585_p17.jpg", "text": "16\nNACA TN 1907\n\narticulated that a change in lag angle would not change the incidence)\nif the pilot failed to alter the pitch manually. The instantaneous\npitch change could be approached by a rotor having articulated blades\nwith no damping about the drag hinge. In this case, the change in lag\nangle would be very rapid, and any attendant change of incidence could\nbe considered practically instantaneous. The intermediate, exponential\ntypes of pitch change could each occur, for articulated blades, with an\ninfinite number of combinations of pilot reaction and degrees of damping\nabout the drag hinge. The more gradual changes are associated with\nslower pilot reaction and greater damping about the drag hinge. For\nrigid blades, the curves represent only different pilot reactions.\n\nVariations of Flapping Angle $\\beta$ with Time After\n\nPower Failure\n\nThe influence of the different rates of pitch change, and of blade\nmoment of inertia, on the variation of flapping angle $\\beta$ with time is\ngiven in figure 5. These calculations were made by the method of\nequation (11), using the induced-velocity decay coefficient $k$ equal\nto unity.\n\nThe variations of $\\beta(t)$ shown in figure 5 are of passing interest\nonly, since the effects of blade flapping are of no importance in the\nvariations of $V(t)$ and $\\Omega(t)$. In spite of the linearizing assumption\nin the solution for $\\beta(t)$ that $\\Omega$ is constant, figure 5 probably\ngives a good indication of stop-settings required on the flapping hinge\nto allow the blades complete freedom in this maneuver. The final values\nof $\\beta$ are different for the case of no pitch change and the cases of\npitch reduction because the final values of $\\Omega$ are not the same.\n\nEffect of Blade Flapping on $V(t)$ and $\\Omega(t)$\n\nComparison (fig. 1) of two computations of $V(t)$ and $\\Omega(t)$, for\ninstantaneous pitch change, shows that the effect of blade flapping is\nnegligible. The case of instantaneous pitch change was chosen for this\ncomparison, since the variations of $\\beta$ are greatest for this\ncase (fig. 5). If the flapping hinge is so directed that a change\nin $\\beta$ causes a change in incidence, then in equation (2) for $\\theta(t)$\nthe influence of flapping must be considered. The value of $k$ was\ntaken to be unity.", "timestamp": "2026-07-22T04:56:35.156379+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 17, "total_pages": 58, "image_filename": "19930082617_p17.jpg", "text": "```markdown\n16\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n6 216.0 X $10^3$\n\n3.86\"\nBand B\nA\nA\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n12 8 4 0 4 8 12 16 20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6\n\nNACA\n\nFigure 5.- Strain diagram of cylinder 73. Band B.\n```", "timestamp": "2026-07-22T04:56:35.525667+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930085471_p9.jpg", "text": "NACA RM No. L8J11\nCONFIDENTIAL\n7\n\nThe data further indicate that no harmful effect on the flutter speed ensues when the center of gravity of the tip weights and the wing coincide. It is observed that the frequency ratio varies from 0.55 to 0.85 and that the torsional frequency has been reduced to as low as one-third the bare wing value. For the largest mass moment of inertia on the wing tip (last case in table II) it was necessary to move the center of gravity farther toward the trailing edge to produce flutter.\n\nIn figure 9, the curves represent theoretical flutter-speed coefficients for one-degree-of-freedom flutter according to reference 1. For the sake of possible interest, the experimental flutter-speed coefficients corresponding, however, to the coupled bending-torsion values are shown superposed on the figure.\n\nIt was hoped that some systematic aspect-ratio effects could be obtained from the present tests, but the results were rather contradictory. Some models were used which spanned the tunnel (except for $\\frac{1}{16}$-inch tip clearance) so that presumably two-dimensional flow over the wing could be expected. By retracting the tip from the boundary layer, flutter of full-span models of NACA 16-010 section could be stopped; however, by retracting the tip from the boundary layer, the flutter amplitude of 3-percent double-wedge models increased. The effect of the subsonic boundary layer at the tip of the models is not known. In one particular case the model fluttered at 99 percent of the theoretical frequency on entering the tunnel and the frequency changed to 128 percent at a smaller amplitude when the model spanned the tunnel. As the model was retracted, the flutter frequency dropped to 99 percent of the theoretical value and fluttered to destruction. A more systematic investigation of the aspect ratio, and tip and shape effects is desirable to clarify various aspects of the problem.\n\nCONCLUSIONS\n\n1. There is rather satisfactory agreement between experimental and calculated flutter-speed coefficients. In general, the theoretical values are conservative.\n\n2. No very pronounced effect of airfoil-section shape on the flutter characteristics was found in these preliminary experiments; however, significant divergence effects were observed on thick blunt-nosed airfoils.\n\n3. The data suggest that for cantilever wings the bending degree of freedom may suppress the one degree of torsional flutter and that coupled bending-torsion effects occur.\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T04:56:35.793323+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 32, "total_pages": 33, "image_filename": "19930082487_p32.jpg", "text": "30\nNACA TN No. 1813\n\n2.2\n2.0\nPrandtl-Meyer\n1.8\n1/2 Prandtl-Meyer\n1.6\nLocal Mach number, M\n1.4\n1.2\n1.0\n.8\n.6\n0\n.1\n.2\n.3\nChordwise station, x/c\nM_o\nO 0.702\n$\\Delta$ .755\n$\\square$ .807\nNACA\n\nFigure 10.- Local Mach number distribution over upper surface of NACA 23015 airfoil section; $\\alpha$, 2°.", "timestamp": "2026-07-22T04:56:38.061110+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 26, "total_pages": 99, "image_filename": "19930082511_p26.jpg", "text": "24\nNACA TN No. 1826\n\nLimiting case of completely open tunnel.- With increasing distance of the vortex from the closed entrance (that is, with increasing $y_1$), expression (9) approaches $-\\frac{i\\Gamma'}{2}$. As was pointed out in part I, the image system normally used to satisfy the boundary condition on an infinitely long, open, two-dimensional tunnel produces no induced flow at the vortex itself, and only after introduction of the additional condition that the upstream flow be horizontal is this value of $-\\frac{i\\Gamma'}{2}$ for the induced-flow correction obtained. In the present development, however, it is seen that the condition of continuity at the entrance lips automatically takes care of this condition on the upstream flow direction, even when the entrance and the vortex are infinitely far apart.\n\nNo further discussion of the completely open tunnel will be given here inasmuch as this case has been adequately treated by the method of images. (See reference 6, p. 302.)\n\nCASE 2 - TUNNEL WITH ONE FIXED EXIT BOUNDARY\n\nPerturbation velocity.- As before, the transformation $z = e^{\\pi\\zeta}$ transforms the tunnel, considered as an infinite strip of unit height, into the upper half of the z-plane. The correspondence between points is shown in figure 14. The conditions on the complex velocity $Q(z)$ are:\n\n(1) On the real axis, I.P.$Q(z) = 0$ for $-1 < z < 1$ and for $z > a$\n(2) On the real axis, R.P.$Q(z) = 0$ for $z < -1$ and for $1 < z < a$\n(3) For $z = \\pm 1$, $Q(z) = 0$\n(4) $Q(z)$ is finite everywhere in the upper half-plane except at $z = a$ and at $z = z_1$. As noted in part I, $Q(z)$ will be infinite at $z = a$\n\nThe function $G(z)$, given by equation (2), will again be used as a factor that is real along the entire real axis and has the desired type of singularity at $z_1$. The functions $\\sqrt{\\frac{1-z^2}{a-z}}$ and $z\\sqrt{\\frac{1-z^2}{a-z}}$ change from pure real to pure imaginary as $z$ passes through $\\pm 1$ or through $a$, and, furthermore, are zero at $z = \\pm 1$ and infinite at $z = a$. At infinity they are of order $z^{1/2}$ and $z^{3/2}$, respectively. Therefore, as before, a linear combination of $G(z)\\sqrt{\\frac{1-z^2}{a-z}}$ and $G(z)z\\sqrt{\\frac{1-z^2}{a-z}}$", "timestamp": "2026-07-22T04:56:45.254989+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 11, "total_pages": 14, "image_filename": "19930082712_p11.jpg", "text": "NACA TN 1998\n\n9\n\nTABLE I\n\nORDINATES FOR THE\n\nNACA 8-H-12 AIRFOIL SECTION\n\n[Stations and ordinates given in percent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| --- | --- | --- | --- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .117 | .2212 | 1.229 | .1535 |\n| .853 | .1474 | -.819 |\n| .358 | .2726 | 1.520 | .2855 |\n| 1.112 | .1703 | -.916 |\n| .6472 | .301 | .3108 | 2.006 |\n| .3052 | 1.696 | .2230 | -1.128 |\n| .3541 | .980 | .2738 | 2.911 |\n| .5432 | 3.020 | .2547 | -1.115 |\n| .41174 | .60368 | .79632 | 1.214 |\n| .77614 | .312 | .4036 | 5.576 |\n| .3125 | -1.736 |\n| .75320 | 1.24452 | 6.911 | .3684 |\n| 5.380 | 1.5555 | 8.086 | .3450 |\n| -1.920 |\n| .787692 | 1.69689 | 9.127 | .42734 |\n| 6.263 | 1.703 | 10.573 | .3706 |\n| -2.059 |\n| 1.067612 | 1.60940 | 11.197 | .43720 |\n| 6.626 | 2.794 | 15.503 | .4436 |\n| -2.212 |\n| 1.204703 | 1.52970 | 19.607 | 1.5489 |\n| 8.605 | 3.670 | 20.393 | .4232 |\n| -2.351 |\n| 1.294074 | 1.43572 | 21.751 | 1.63799 |\n| 9.213 | 4.544 | 25.216 | .4251 |\n| -2.117 |\n| 1.336678 | 1.39441 | 29.969 | 1.71594 |\n| 9.533 | 5.4036 | 30.031 | .4419 |\n| -2.155 |\n| 1.320176 | 1.33122 | 35.171 | 1.69776 |\n| 9.152 | 6.268 | 34.826 | .4482 |\n| -2.190 |\n| 1.261677 | 1.25254 | 40.292 | 1.6254 |\n| 9.030 | 7.1494 | 39.708 | .4489 |\n| -2.191 |\n| 1.178808 | 1.1698 | 45.360 | 1.5156 |\n| 8.1120 | 8.030 | 44.610 | .4457 |\n| -2.176 |\n| 1.07321 | .9070 | 50.390 | 1.37988 |\n| 7.666 | 8.9241 | 49.610 | .4385 |\n| -2.136 |\n| 0.951313 | .96862 | 55.387 | 1.2231 |\n| 6.795 | 9.830 | 54.613 | .4279 |\n| -2.377 |\n| .81816 | .82104 | 60.358 | 1.05228 |\n| 5.816 | 10.735 | 59.612 | .4122 |\n| -2.290 |\n| .67905 | .7155 | 65.311 | .8730 |\n| 4.850 | 11.644 | 64.689 | .3920 |\n| -2.178 |\n| .537321 | .615 | 70.250 | .6984 |\n| 3.838 | 12.556 | 69.750 | .3641 |\n| -2.031 |\n| .397321 | .5331 | 75.181 | .51064 |\n| 2.838 | 13.466 | 74.816 | .3348 |\n| -1.860 |\n| .265301 | .4521 | 80.118 | .3411 |\n| 1.895 | 14.378 | 79.882 | .2961 |\n| -1.615 |\n| .146661 | .3706 | 85.063 | .18828 |\n| 1.016 | 15.288 | 84.940 | .2491 |\n| -1.381 |\n| .054702 | .2620 | 90.016 | .06179 |\n| .315 | 16.199 | 89.981 | .18710 |\n| -1.051 |\n| .016661 | .1709 | 94.999 | -.02192 |\n| -.119 | 17.110 | 95.005 | .1132 |\n| -.629 |\n| 18.0100 | .000 | 0 | 18.0100 | .000 |\n| 0 |\n\nL.E. radius: 1.325 .1855\n\nSlope of radius through L.E.: 0.314\n\n.2385\n\n[Figure: Airfoil section plot with x/c axis from 0 to 1.0]\n\nNACA", "timestamp": "2026-07-22T04:56:48.431747+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 31, "total_pages": 41, "image_filename": "19930082476_p31.jpg", "text": "NACA TN No. 1801\n29\n\n[Figure: Photograph of the model spinning in the Langley 20-foot free-spinning tunnel. The image shows a large indoor facility with multiple windows and people observing. A label on the bottom right reads \"NACA L-50738\".]\n\nFigure 3.- Photograph of the model spinning in the Langley 20-foot free-spinning tunnel.", "timestamp": "2026-07-22T04:56:50.196473+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 16, "total_pages": 36, "image_filename": "19930082614_p16.jpg", "text": "14 NACA TN 1939\n\nbe used with the graphs is the component measured normal to the flight path in a vertical plane tangent to the flight path.\n\nThe procedure in the calculation of the speed changes within each time increment is the same with a variable flight-path angle as with a constant angle. The flight-path angle and the altitude are progressively evaluated for the beginning of each time interval and their values are estimated for the interval by the method just described. The acceleration of the end of the increment $\\Delta t$ is obtained graphically (fig. 1) and the velocity is computed. Step-by-step calculations furnish the velocity variation during the complete interval being investigated.\n\nDRAG CHARACTERISTICS OF AERODYNAMIC BRAKES\n\nThe characteristics of various aerodynamic brakes have been investigated in wind-tunnel and flight tests. (See references 3 through 7.) Results of such tests are summarized in this report. Data are presented in the form of increments in drag coefficient which are attributable to the brakes. It is to be expected that in some cases such increments are affected by the particular location of the brakes on the wing, the fuselage, or elsewhere, and by their proximity to other components of the airplane. The air brakes shown are representative of installations in which air brakes are added to typical fuselages or at different locations on wings.\n\nGeometric data and incremental drag coefficients for the aerodynamic brakes shown in figure 6 are presented in table I. These data indicate increments in the drag of the airplane which are small in comparison with the values required in a steep dive for an airplane having a moderate wing loading (fig. 4). Increased drag can be obtained by increasing the relative size of the brake. However, increasing the size is practical only within the limitations of available space into which the brake may be retracted. The size is limited also by considerations of weight of the structure transmitting the aerodynamic loads, and by the effects of large brakes upon the trim, stability, and buffeting of the airplane.\n\nTable I indicates that the drag coefficients (based upon the areas of the air brakes) vary over a wide range, depending upon the shape and location of the brake on the airplane. The lowest drag coefficient was measured for the picket-fence type of brake (type N). This low value of brake drag coefficient might be expected since the brake area used as a reference is more than twice the actual frontal area. The highest drag resulted from solid brakes at forward positions on the wing (types F and G). This forward location of the brakes results in a spoiling action which causes changes in the lift as well as drag. Since the drag increments in table I are for zero lift, a part of the", "timestamp": "2026-07-22T04:56:52.190709+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 25, "total_pages": 50, "image_filename": "19930082496_p25.jpg", "text": "24\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T04:56:52.390782+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 4, "total_pages": 20, "image_filename": "19930085536_p4.jpg", "text": "2\nNACA RM No. E8K05\n\nAn experimental investigation was undertaken at the NACA Lewis\nlaboratory to check theoretical calculations for a conical body of\nelliptic cross section. The results are compared with calculations\nbased on the linearized theory given in reference 6.\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| | |\n| :--- | :--- |\n| $C_p$ | pressure coefficient |\n| $K$ | constant proportional to source strength |\n| $M$ | Mach number |\n| $m$ | slope of line source with respect to x-axis |\n| $U$ | free-stream velocity |\n| $U_r$ | radial perturbation-velocity component (cylindrical coordinate) |\n| $U_x$ | axial perturbation-velocity component |\n| $U_x'$ | perturbation-velocity component parallel to free-stream direction |\n| $U_\\theta$ | tangential perturbation-velocity component (cylindrical coordinate) |\n| $x,r,\\theta$ | cylindrical coordinates |\n| $\\alpha$ | angle of attack, degrees |\n| $\\beta$ | cotangent of Mach angle, $\\sqrt{M^2-1}$ |\n| $\\gamma$ | ratio of specific heats |\n| $\\delta$ | angular position of line source measured from $\\theta = \\pi/2$ plane |\n| $\\Psi$ | angle of yaw, degrees |", "timestamp": "2026-07-22T04:56:54.615262+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 12, "total_pages": 66, "image_filename": "19930082914_p12.jpg", "text": "NACA TN No. 1857\n\nnearly parallel as possible by eye, and then leaving them there, because it is possible to produce all fringe orientations by adjustment of the splitter plates alone (except for adjustment for white-light fringes which is made with the compensating plates). Then the plates are set nearly parallel to the mirrors, by simple inspection. A small light source is placed 15 feet or more away from the interferometer, and the light is directed at plate $S_1$. Two sets of cross hairs are set up, one near the light source and the other near plate $S_1$. Two screens are set up on the opposite side of the interferometer in such positions that a lens placed after plate $S_2$ focuses one set of cross hairs on one screen and the other set on the other screen. Each set of cross hairs produces two images on one screen. One image is produced by the light that is transmitted through the lower part of the interferometer, and the other image by the light that goes through the upper part. The splitter plates are then so adjusted that the two images of one set of cross hairs coincide, or merge into a single image. This procedure is repeated for the other set of cross hairs. As the adjustments for the two sets of images are not independent, they must be continued until the images of both sets of cross hairs appear to be single. The light source is now replaced by a monochromatic light source. This light source should be sufficiently close to being monochromatic that many hundreds of fringes can be produced. The reason for this is that at this stage of the adjustment the two optical paths through the interferometer may be considerably different. If such is the case, and a light source that would give relatively few fringes is used, then the fringes cannot be found because their apparent location would be above or below the test section and out of the light beam. A satisfactory light source is a sodium-arc lamp, or a General Electric B-H6 mercury lamp operated at an undervoltage of about 60 volts. The lens in the beam emergent from the interferometer is replaced by a telescope placed some 15 feet beyond the interferometer. Interference fringes generally can then be brought into focus at some point between the light source and the telescope. If they cannot be found, either the elimination of double cross-hair images has not been sufficiently achieved or the difference in optical-path length is too great. Inspection of each cross-hair image (on the screens) with a magnifying glass will usually reveal that the elimination of doubling has not been completely effected.\n\nOnce the fringes have been located, the monochromator is set up. A B-H6 mercury lamp is placed at the entrance to the monochromator. The slits are narrowed to about 0.3 millimeter and the prism set to pass the 5460-angstrom line. The next adjustment is to tilt the fringes into the desired orientation, say a horizontal position, by rotating a splitter plate.\n\nThe next two adjustments are to move the fringes into the test section and to adjust them to the desired width, or spacing. These two adjustments are made simultaneously by alternate rotation of the two splitter plates and by observing through the telescope the location and the spacing of the fringes. This adjustment is not difficult. For example, suppose", "timestamp": "2026-07-22T04:56:57.560530+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 39, "total_pages": 47, "image_filename": "19930093773_p39.jpg", "text": "38\nNACA RM E9G09\n\nCorrected specific fuel consumption based on net thrust, $W_f/F_n\\sqrt{\\theta}$\nlb/(hr)/(lb thrust)\n\nAltitude\n(ft)\nO 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n◁ 50,000\n\nCorrected engine speed, $N/\\sqrt{\\theta}$, rpm\n\n(d) Specific fuel consumption.\nFigure 6. - Continued. Effect of altitude on variation of corrected\nengine performance with corrected engine speed at flight Mach\nnumber of 0.21.", "timestamp": "2026-07-22T04:57:12.795764+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 34, "total_pages": 62, "image_filename": "19930082485_p34.jpg", "text": "NACA TN NO. 1810\n\nTubes to pressure\ntaps on blade\nsurface\n\nStatic-pressure\ntaps\n\nTotal-pressure\ntaps\n\nBlades\n\nOuter\nshroud\n\nAir flow\n\nInner shroud\n\nSurvey probe\n\nProtractor\n\nTraversing\nmechanism\n\nNACA\n\nFigure 3.- Turbine-blade-cascade experimental equipment.\n\n33", "timestamp": "2026-07-22T04:57:15.727527+00:00"}